1. Field of the Invention
This invention resides in the field of combustion systems for liquid-fuel rocket engines, and is specifically concerned with heat exchange structures for expander cycle rocket engines.
2. Description of the Prior Art
Rocket engines such as those used on space missions require both high thrust and a durable construction that can withstand the extreme conditions of temperature and pressure encountered upon takeoff and in flight and that will enable repeated use of the engines in successive firings. Expander cycle rocket engines use regenerative cooling to achieve both high thrust and durability. In the typical expander cycle rocket engine, the internal walls of the combustion chamber and nozzle are cooled by uncombusted fuel fed by a turbopump into a jacket that surrounds the chamber and nozzle. The heated fuel emerging from the jacket is cycled through the turbine side of the turbopump to serve as the driving medium for the pump. The expanded fuel emerging from the turbine side of the pump then passes into the combustion chamber for combustion with oxidizer. The cycle thus cools the combustion chamber while simultaneously converting a portion of the heat generated by the engine to higher flow rates of fuel and oxidizer to, and hence a higher pressure in, the combustion chamber. The remainder of the heat is retained by the fuel as it enters the combustion chamber, thus preheating the fuel.
The amount of power that the turbopump can extract from the expander cycle to generate pressure in the combustion chamber and the amount of preheating of the fuel before it enters the chamber are limited by the amount of energy that can be extracted from the engine through the cooling jacket. The maximum chamber pressure that has been demonstrated by expander cycles of the prior art is thus typically about 500 psia. Attempts to achieve higher pressures have included the use of heat exchange tubing with bumps on the inner and outer tubing surfaces to produce turbulence at these surfaces. This has met with limited success, increasing the heat transfer rate by only 20% to 40%. Much higher rates of heat transfer are needed if the full potential of the expander cycle is to be realized. Other attempts have involved adding chamber length to increase the surface area and residence time in the jacket. This however adds substantially to the weight of the chamber and to the axial dimension of the rocket engine.
The most efficient heat exchange structures are those that provide the most intimate contact between the medium to be cooled and the coolant, i.e., the largest surface area, the thinnest separating walls, and the narrowest flow channels. One type of structure that offers both of these features is a laminated platelet stack formed by the bonding together of very thin sheets of heat-conductive metal, the sheets having been individually etched prior to bonding to form very narrow, and in some cases intricate, flow passages. Platelet stacks such as these have therefore been used as construction panels for coolant jackets, with the coolant flowing through these very narrow passages. Heat transfer efficiency is still limited however by the fact that the heat extracted from the combustion gas in the chamber is drawn only from the boundary layer.
The present invention resides in a novel rocket nozzle structure that provides increased heat transfer efficiency by using the pressure drop arising from flow through the combustion chamber to divert portions of the combustion gas into channels in the chamber wall where the combustion gas is in intimate heat exchange contact on two sides with the fuel being preheated. The terms xe2x80x9ccombustion chamberxe2x80x9d and xe2x80x9cchamberxe2x80x9d are used in this specification and the appended claims to denote the portion of the rocket nozzle upstream of the throat.
Heat exchange in the chamber between the combustion gas and the uncombusted fuel in this invention thus occurs by heat fluxes in two opposing directions, rather than limiting the heat transfer to a single direction at the internal surface of the chamber wall. This augmented heat exchange is achieved by layers of internal channels in the chamber wall, the combustion gas channels occupying at least one layer and the uncombusted fuel channels occupying layers positioned immediately adjacent to, and on opposing sides of, each layer of combustion gas channels. (The channels carrying uncombusted fuel are also referred to herein as xe2x80x9ccoolant channelsxe2x80x9d due to the cooling effect they have on the chamber.) The innermost layer is thus a layer of coolant channels drawing heat on one side from the boundary layer in the chamber interior and on the other side from the adjacent layer of channels carrying combustion gas that has been diverted from the chamber interior. Likewise, the layer of channels carrying combustion gas that has been diverted from the chamber interior heats both the innermost layer of coolant channels and a layer of coolant channels on the opposite side, i.e, the third layer from the chamber interior. Similar two-directional heat exchange occurs for every layer of channels except the outermost layer that is closest to the outer surface of the chamber wall.
The combustion gas channels are of limited length, drawing combustion gas from and returning it to the chamber interior, each channel aligned generally in the axial direction or in a direction having an axial component. Each channel has two ports opening into the chamber, the two ports being displaced from each other in the direction of flow in the chamber, one port thus being upstream of the other. Because of the pressure gradient that spontaneously occurs in the channel during the flow of the combustion gases, the chamber pressure at the upstream port is higher than the chamber pressure at the downstream port, causing combustion gas to enter the channel at the upstream port, flow through the channel and exit the channel at the downstream port. In preferred embodiments of the invention, the length of each channel is considerably less than the length of the chamber. The combustion gas channels are distributed around the circumference of the nozzle, and increased efficiency is obtained by including two or more rows of combustion gas channels successively positioned axially along the chamber wall. The boundary layer is thus disrupted at each row of ports, allowing each row of channels to draw fresh and relatively hot combustion gas from the chamber interior for heat exchange with the uncombusted fuel in the coolant channels.
The coolant channels are supplied by a source of fresh fuel, which in the case of expander cycle rocket engines is pumped to the coolant channels by the turbopump. The coolant channels preferably extend the full length of the combustion chamber, and discharge into a common line that leads back to the turbine side of the turbopump. The entries to and exits from the coolant channels are preferably arranged such that the flow of uncombusted fuel through the coolant channels is countercurrent to the flow of combustion gas through the chamber and through the combustion gas channels.
These and other features of the invention, as well as various preferred embodiments, are described in greater detail below.